The invention relates to the field of gas turbine engines, and more specifically, to an improved gas turbine engine using compressor fluid to maintain a thermal boundary layer between turbine blades and heated fluid from the combustor.
A type of prior art gas turbine has a compressor, a fuel source, a source of air for combustion, a casing, and a combustor to prepare heated fluid from fuel and combustion air. The combustor has a combustion zone that is connected to the fuel source and the source for combustion air. It contains a cooling zone for cooling the resulting heated fluid before reaching the turbine. The combustor cooling zone is connected to the compressor. The heated fluid temperature fluctuates depending on operating conditions. In conventional gas turbine engines, these temperature fluctuations result in strong temperature-induced stresses imparted on the engine components.
In these prior art gas turbines, virtually the entire compressor fluid flow is directed to the combustor. Fluid heated in the combustor is cooled by the compressor fluid flow in the combustor cooling zone. This engine has a turbine rotor disk with blades that receive heated fluid from the combustor. The temperature of this heated fluid is quite high and, under certain conditions, the fluid can overheat the turbine rotor disk blades. To prevent such overheating, each blade has an interior channel that receives air from the compressor. As a result, part of the fluid coming from the compressor (about 3% to 5% of the total flow) is supplied to the interior channel of the blades to keep their temperature within design limits. One example of this gas turbine engine is disclosed in U.S. Pat. No. 3,826,084 to Branstrom. et al.
The heated fluid also has to be cooled after fuel combustion. Normally, this is done in the combustor, to which the major part of the fluid from the compressor would be admitted.
Thus, in this prior art gas turbine, substantially all of fluid coming from the compressor is supplied to the combustor cooling zone to cool the fluid before it enters the turbine. When fluid from the compressor is mixed with the heated fluid in the combustor cooling zone, about 3% to 5% of the fluid""s energy is lost. Diverting about 3% of fluid from the compressor to the turbine rotor disk blades results in another 3% in losses. In addition, the combustor for this prior art gas turbine has to be made larger to accommodate the cooling zone.
It is an object of the invention to provide a gas turbine engine that would have a better efficiency and lower harmful emissions.
It is another object of the invention to reduce the cooling zone losses of the prior art combustor.
Another object of the invention is to provide a more compact gas turbine engine.
A further object of the invention is to prolong the life of the gas turbine engine.
The foregoing objects are accomplished through the design of a gas turbine engine having a turbine blade flow portion located between a blade inlet edge and outlet edge, and having a compressor to supply a fluid flow, comprising a substantial portion of compressor fluid output, along the outside surface of the blade flow portion, thus providing a thermal insulating boundary layer over which flows heated fluid from the combustor. Heated fluid formed in the combustor flows without further cooling directly to the turbine blades, thereby reducing losses associated with cooling.